The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in several rows or stages of turbine blades which power the compressor, and power a fan in a turbofan aircraft engine application, or power an external drive shaft in marine and industrial applications.
The high pressure turbine (HPT) includes one or more stages which first receive the hot gases from the combustor, and are typically internally cooling by channeling therethrough a portion of pressurized air bled from the compressor. Each turbine blade includes an airfoil extending radially outwardly from an integral platform and supporting dovetail. The airfoil is hollow and includes various cooling circuits therein having inlets at the base of the dovetail for receiving the bleed or cooling air.
Turbine blade cooling is crowded with various configurations of the cooling circuits therein tailored for accommodating the varying heat loads from the combustion gases over the opposite pressure and suction sides of the airfoil and between the axially opposite leading and trailing edges and the radially opposite inner root and outer tip. Each airfoil typically includes rows of film cooling holes through the sidewalls thereof which discharge the spent cooling air in corresponding thermally insulating films over the external surface of the airfoil.
The internal cooling circuits include radial channels or legs axially separated from each other by corresponding partitions or bridges which extend radially along the span of the airfoil. Dedicated cooling legs may be used directly behind the leading edge and directly in front of the trailing edge for specialized cooling thereof. The midchord region of the airfoil may also include dedicated cooling channels typically in the form of serpentine circuits having multiple radial legs which alternate the radial flow of the cooling air between the root and tip of the airfoil.
One or more serpentine cooling circuits may be used in each airfoil either combined with or independent from the cooling legs along the leading and trailing edges. The cooling circuits may also include various forms of short ribs or turbulators along the inner surface of the pressure and suction sidewalls for tripping the cooling air to increase the heat transfer thereof.
Turbine blades found in gas turbine engines are subject to the local environment in which the engine is operated. And, minimum weight of the engine is typically a paramount design objective, particularly for aircraft engines, which limits the size and complexity of the engine and associated equipment when integrated into the aircraft.
For example, operating a gas turbine engine in a dusty environment will carry dust particles through the compressor, which dust particles are also entrained in the cooling air used for the turbine blades. The cooling circuits of the turbine blades are relatively small, with small features therein, including the various rows of film cooling holes. A typical film cooling hole is about 10–15 mils in diameter and is susceptible to dust accumulation during operation.
Accordingly, it is common practice to include relatively large dust holes in the airfoil tip to permit entrained dust particles to readily exit the airfoil and minimize dust accumulation therein or in the smaller film cooling holes. A typical dust hole is about 25 to 60 mils in diameter which is several times the diameter of the small film cooling holes.
Furthermore, the tip region of the typical turbine blade also requires dedicated cooling to ensure its durability and long useful life. The typical airfoil tip is a thin flat plate which closes the radially outer end of the airfoil, and typically includes short extensions of the pressure and suction sidewalls in the form of squealer or tip ribs. The squealer ribs define an outwardly open tip cavity in which the dust holes are located for discharging the cooling air and any dust from the airfoil. The tip also typically includes several smaller cooling holes for cooling the tip itself and the adjacent squealer ribs.
Since the dust holes are typically much larger than the film cooling holes and the tip cooling holes, they themselves have little cooling efficacy and correspondingly increase the flowrate of the cooling air which must be channeled through each airfoil. Any such excess air channeled through the airfoil correspondingly decreases the overall efficiency of the gas turbine engine since the bleed air is not used in the combustion process.
Compounding this problem of the large dust holes is the typical need for multiple large dust holes in an individual turbine blade for multiple independent cooling circuits used therein. Since the pressure distribution of the combustion gases varies over the external surface of the airfoil between the leading and trailing edges and along the pressure and suction sides of the airfoil, independent cooling circuits are typically provided inside each airfoil and fed with a common pressure bleed air.
As the cooling air is channeled through the independent circuits in the airfoil it experiences pressure losses or drops between the inlet of each cooling circuit and the various outlets thereof, such as the film cooling holes along the airfoil sidewalls, trailing edge holes along the thin trailing edge, and the tip cooling holes in the airfoil tip. Each circuit and its various outlet holes are specifically designed for maintaining a suitable backflow margin at each of the outlet holes to prevent ingestion of the hot combustion gases into the airfoil during operation. Backflow margin is a primary design objective in configuring the various cooling circuits inside each turbine blade.
Accordingly, it is desired to provide an improved turbine blade for minimizing air discharge through the dust holes.